German Patent Publication 3,535,779 discloses a thrust nozzle for a high power engine, for example for a booster rocket or a reusable spacecraft, which has a rotationally symmetric contour. The circular cross-section of the known thrust nozzle tapers or becomes narrower starting from the combustion chamber in a direction toward the constricted throat of the nozzle, whereafter it again expands or becomes wider outwardly. Such a known rotationally symmetric contour is simple to manufacture using known production technology and makes it possible to effectively withstand the arising gas pressure forces.
Due to the high operating temperatures of approximately 3000.degree. C., it is in any event necessary that the thrust nozzle must be effectively cooled. In the conventional thrust nozzle, which comprises an inner casing or jacket made of a copper alloy, the necessary cooling is achieved by providing cooling channels extending in the circumferential direction or in the axial direction in the inner jacket. A cooling medium, such as the liquid hydrogen that is to be combusted in the combustion chamber or thrust nozzle, flows through the provided cooling channels and thereby cools the thrust nozzle. The conventional thrust nozzle further includes an outer support casing or jacket that externally surrounds the inner jacket in a joint-free manner, whereby this outer support jacket serves to take up and withstand the forces arising from the combustion gas pressure. For this purpose, the support jacket must comprise a high tensile strength, while its thermal resistance or temperature stability is of lesser importance due to the cooling that is provided.
Efforts are currently being pursued for developing hypersonic aircraft, which will similarly utilize such thrust nozzles. Due to weight, fuel capacity, and cost considerations, these thrust nozzles must achieve a high efficiency of thrust generation, whereby several engines are typically to be arranged clustered adjacent one another. In order to meet these requirements, there have already been proposals of thrust nozzles of which the cross-sectional contour transitions from a round cross-section in the area of the combustion chamber to a quadrilateral cross-section in the area of the exhaust outlet of the thrust nozzle or even in the area of the constricted throat of the thrust nozzle. Such a configuration in turn requires that the thrust nozzle wall has a complex curved contour.
On the one hand, the relatively soft inner jacket must have an inner contour exactly matching a prescribed shape in order to achieve an optimal combustion and gas through-flow. On the other hand, the outer supporting jacket must be sufficiently rigid and configurationally stiff for strength reasons. Therefore it becomes impossible to achieve a matching or adaptation of the outer jacket to the form of the inner jacket. As a result, it becomes necessary to manufacture the two jackets with a very high degree of matching shape accuracy, which necessitates the use of very complicated and costly production processes in view of the complicated geometry that is involved. A further disadvantage is that it is possible for hollow spaces or gaps to remain between the two jackets after they are joined together, whereupon these hollow spaces or gaps can lead to deformations, cracking and ultimately failure of the inner jacket during operation of the combustion chamber and thrust nozzle. As a result, the operating life of the rocket engine is limited.
To overcome the above discussed disadvantages, German Patent Publication 4,015,204 discloses a thrust nozzle for an aeronautic or aerospace engine comprising an inner casing or jacket that has a high thermal conductivity and that is provided with a plurality of cooling channels, a rigid and strong support casing or jacket externally surrounding the inner jacket, and an intermediate layer that is integrally cast in place between the inner jacket and the outer support jacket. In this manner, the intermediate layer serves to compensate for any production tolerance between the inner jacket and the outer support jacket so that the requirements of shape accuracy of the two jackets can be reduced.
German Patent Publication 4,115,403 discloses a thrust nozzle wall construction for expansion ramps and hot gas nozzles that comprises an outer supporting structure arranged facing away from the hot gas, as well as a multi-layered inner structure having a plurality of mutually spaced-apart cooling channels that run along the inner structure facing toward the hot gas that is present inside the nozzle. In order to achieve a high thrust output, and in order to simplify the exchangeability of the nozzles between different types of engines, thrust nozzles having a quadrilateral configuration are particularly suitable. However, the nozzle walls of such thrust nozzles are subjected to high pressure forces and high temperatures. In contrast to the walls of circular cross-section nozzles, flat planar nozzle walls of quadrilateral nozzles or combustion chambers are subjected to high bending moments due to the pressure forces of the combustion gases. As a result, bulging or bending deformations and overstrain conditions can arise in the thrust nozzle walls, which jeopardize the proper functioning of the thrust nozzle according to specifications. An additional difficulty arises in view of the so-called bi-metal effect, due to the different metal components of the inner jacket and of the outer jacket and due to the temperature differences through the multi-layered wall. In order to avoid thrust losses and leakage flows, it is therefore necessary to provide very rigid, form-stable, cooled walls for the combustion chamber and the thrust nozzle.
For the above reason, this known thrust nozzle wall comprises an inner structure that consists of an inner thermal conduction layer that is impinged upon by the hot gas, as well as a heat resistant gliding layer. The required cooling channels are embedded in the thermal conduction layer, and these are elastically connected to the supporting structure by means of a plurality of holding elements that penetrate through the gliding layer. In this context, the gliding layer can consist of a ceramic granulate, while the thermal conduction layer consists of copper. The holding elements can be embodied in the form of small pipes or tubes. However, due to the required minimum strength in this context, an adequate straining distance is not available when the thrust nozzle is subjected to the extremely high thermal loading that is typical in the operation of high power engines. Due to the high thermally induced tensions, which cause considerable plastic strain deformation of these holding elements particularly, the operating life of the wall construction is drastically limited.
The above discussed limitation of the operating life is predominantly caused by a failure or defect formation, for example the formation of a crack in the combustion chamber wall, after a limited number of load cycles and associated plastic deformations and creeping of the components due to the thermally limited strains, in other words due to the secondary stresses caused by the high thermally induced stresses amounting to approximately 80% of the total load.
For the above reasons, not only is the reusability of the combustion chamber and thrust nozzle or overall engine system sharply limited, but also the total costs of the booster or other propulsion system are significantly increased. Moreover, thrust losses and overloading of the various engine components, including the known turbo-pumps for pumping the fuel, oxidizer, and/or coolant, also arise during operation of the engine, and, for example, lead to the formation of cracks or the like in these various components.
Even in the case if intermediate materials are arranged between the inner wall that is exposed to the hot gases and the outer supporting structure, such as sintered aluminum or foamed aluminum materials for example, which can take up and effectively absorb relatively high deformations, irreversible deformations in the plastic deformation range ultimately also arise, so that any such constructions are only suitable for single use engine concepts.
The above discussed support elements arranged between the inner hot gas wall and the outer surrounding structure, when made of conventionally known materials, are intended to achieve a defined yielding in the direction of their respective length as a result of associated widthwise or crosswise strain during operation of the high power engine. However, the known materials for making such supporting elements do not comprise an adequate elastic strain characteristic. As a result, the elastic strain required for the desired compensating effects during operation of the engine is not provided.
German Patent Publication 3,136,252 describes a ceramic combustion chamber wall for burner heads and combustion chambers, which are installed in combustion and drive plants or assemblies. In this context, the combustion chamber wall of the combustion space is an integrated component of a ceramic heat exchanger with single-stage or multi-stage flow channels. By means of strand pulling or extrusion techniques, film techniques, and winding or wrapping techniques, it is possible to produce such combustion chamber walls with parallel-flow, counter-flow, and/or cross-flow heat exchangers, as they are otherwise usually produced in metal embodiments. The use of silicon carbide or silicon nitride is suggested as the material for the combustion chamber wall.